The present invention pertains to a method for repairing gas turbine engine airfoil parts. More particularly, the present invention pertains to a method for restoring critical gas path flow area dimensions in cast nickel or cobalt-base superalloy airfoil components of a gas turbine engine.
Airfoil parts, such as blades and vanes, are critical components in the gas turbine engines that are used to power jet aircraft or for the generation of electricity. Each airfoil part is an individual unit having a root or attachment section and an airfoil section. The airfoil section has specific cordal and length dimensions that define the airfoil characteristics of the part. The root section is engaged with and held by a housing member. A plurality of the airfoil parts are thus assembled with the housing member to form a disc or ring. Blades, which during operation are rotating part, are assembled into and disc. Vane, which remain stationary, are assembled into a nozzle or vane ring. In the operating gas turbine engine the assembled rings and discs, determine the path of the intake, combustion and exhaust gasses that flow through the engine.
The airfoil part may be either a rotating component or a non-rotating component of the gas turbine engine. If the part is a rotating component, during operation of the turbine engine the part is subjected to centrifugal forces that exert deforming stresses. These deforming stresses cause creep rupture and fatigue problems that can result in the failure of the part. Non-rotating components, such as vanes, are not subjected to centrifugal forces that exert deforming stresses. However, like the rotating parts, these parts are subjected to other deformation such as from hot gas erosion and/or foreign particle strikes. This deformation results in the alteration of the dimensions of the airfoil section. The alteration of the dimensions of the airfoil section can detrimentally modify the airflow through the gas turbine engine which is critical to the engine's performance.
An example of a non-rotating airfoil part is the 2nd stage vane of the Pratt & Whitney JT8D model 1 through 17R gas turbine engine. This part is manufactured by the "lost wax" or "investment casting" process. The vane is cast from one of several highly alloyed nickel or cobalt-base materials. As a new part in a new gas turbine engine, or as a new spare part in an overhauled engine, it begins its life cycle with a protective diffusion coating on its airfoil surfaces and a wear coating on surfaces known to have excessive wear patterns.
When the gas turbine engine is operating, the vane will see temperatures of about 1500 degree F. Since the vane does not rotate and thus is not subject to creep rupture, its demise is most often influenced by the number of times it is repaired. The reason for this is the repair process itself.
The repair process consists of the following operations:
1.) degrease, wash to remove engine carbon, etc. PA1 2.) grit blast to remove wear coatings, and any sulfidation which is present PA1 3.) chemically remove the diffusion coating PA1 4.) blend to remove nicks, dents, etc. PA1 5.) weld, grind, polish etc.
The repair operations that remove metal by chemical stripping, grit blasting, blending and polishing shorten the life cycle of the vane. The coating removal is a major contributor because it is diffused into the parent metal. When certain minimum airfoil dimensions cannot be met the part is deemed non-repairable and must be retired from service. Thus, there is a need for a method for repairing gas turbine engine airfoil parts that effectively and efficiently restores the airfoil dimensions of the part.
On another front, during the manufacture of metal components a coating operation is performed to provide a coating material layer on the surface of a component substrate. The coating material layer is formed to build-up the metal component to desired finished dimensions and to provide the finished product with various surface attributes. For example, an oxide layer may be formed to provide a smooth, corrosion resistant surface. Also, a wear resistant coating, such as Carbide, Cobalt, or TiN is often formed on cutting tools to provide wear resistance.
Chemical Vapor Deposition is typically used to deposit a thin film wear resistant coating on a cutting tool substrate. For example, to increase the service life of a drill bit, chemical vapor deposition can be used to form a wear resistant coating of Cobalt on a high speed steel (HSS) cutting tool substrate. The bond between the substrate and coating occurs primarily through mechanical adhesion within a narrow bonding interface. During use, the coating at the cutting surface of the cutting tool is subjected to shearing forces resulting in flaking of the coating off the tool substrate. The failure is likely to occur at the narrow bonding interface.
FIG. 12(a) is a side view of a prior art tool bit coated with a wear resistant coating. In this case, the wear resistant coating may be applied by the Chemical Vapor Deposition method so that the entire tool bit substrate receives an even thin film of a relatively hard material, such as Carbide, Cobalt or TiN. Since the coating adheres to the tool bit substrate mostly via a mechanical bond located at a boundary interface, flaking and chipping off the coating off of the substrate is likely to occur during use, limiting the service life of the tool bit. FIG. 12(b) is a side view of a prior art tool bit having a fixed wear resistant cutting tip. In this case, a relatively hard metal cutting tip is fixed to the relatively soft tool bit substrate. The metal cutting tip, which is typically comprised of a Carbide or Cobalt alloy, is fixed to the tool bit substrate by brazing. During extended use the tool bit is likely to fail at the relatively brittle brazed interface between the metal cutting tip and the tool substrate, and again, the useful service life of the tool bit is limited.
Another coating method, known as Conventional Plasma Spray uses a super heated inert gas to generate a plasma. Powder feedstock is introduced and carried to the workpiece by the plasma stream. Conventional plasma spray coating methods deposit the coating material at relatively low velocity, resulting in voids being formed within the coating and in a coating density typically having a porosity of about 5.0%. Again, the bond between the substrate and the coating occurs primarily through mechanical adhesion at a bonding interface, and if the coating is subjected to sufficient shearing forces it will flake off of the workpiece substrate.
Another coating method, known as the Hyper Velocity Oxyfuel (HVOF) plasma thermal spray process is used to produce coatings that are nearly absent of voids. In fact, coatings can be produced nearly 100% dense, with a porosity of less than 0.5%. In HVOF thermal spraying, a fuel gas and oxygen are used to create a combustion flame at 2500 to 3100.degree. C. The combustion takes place at a very high chamber pressure and a supersonic gas stream forces the coating material through a small-diameter barrel at very high particle velocities. The HVOF process results in extremely dense, well-bonded coatings. Typically, HVOF coatings can be formed nearly 100% dense, with a porosity of &lt;0.5%. The high particle velocities obtained using the HVOF process results in relatively better bonding between the coating material and the substrate, as compared with other coating methods such as the Conventional Plasma spray method or the Chemical Vapor Deposition method. However, the HVOF process also forms a bond between the coating material and the substrate that occurs primarily through mechanical adhesion at a bonding interface.
Detonation Gun coating is another method that produces a relatively dense coating. Suspended powder is fed into a long tube along with oxygen and fuel gas. The mixture is ignited in a controlled explosion. High temperature and pressure is thus created to blast particles out of the end of the tube and toward the substrate to be coated.
An example of using HVOF or Detonation Gun coating techniques is disclosed in U.S. Pat. No. 5,584,663, issued to Schell. This reference discloses that the tips of turbine blades can be formed by melting and fusing a powder alloy. Preferably, the blade tip is generated by depositing molten metal alloy powder in multiple passes. Squealers at the perimeter of the blade tip may be formed using methods such as Detonation Gun or HVOF spray methods. The forming step may be used to generate a near- net shaped blade tip, and a subsequent machining step may be employed to generate the final or preferred shape of the blade tip.
Casting is a known method for forming metal components. Typically, a substrate blank is cast to near-finished dimensions. Various machining operations, such as cutting, sanding and polishing are performed on the cast substrate blank to eventually obtain the metal component at desired finished dimensions. A cast metal component will typically have a number of imperfections caused by voids and contaminants in the cast surface structure. The imperfections may be removed by machining away the surface layer of the component, and/or by applying a surface coating.
The manufacture of metal components often entails costly operations to produce products with the desired surface texture, material properties and dimensional tolerances. For example, a known process for manufacturing a metal component requires, among other steps, making a casting of the metal component, treating the metal component using a Hot Isostatic Pressing (HIP) treatment process, and then machining the metal component to remove surface imperfections and obtain the desired dimensional tolerances.
HIP treatment is used in the densification of cast metal components and as a diffusion bonding technique for consolidating powder metals. In the HIP treatment process, a part to be treated is raised to a high temperature and isostatic pressure. Typically, the part is heated to 0.6-0.8 times the melting point of the material comprising the part, and subjected to pressures on the order of 0.2 to 0.5 times the yield strength of the material. Pressurization is achieved by pumping an inert gas, such as Argon, into a pressure vessel. Within the pressure vessel is a high temperature furnace, which heats the gas to the desired temperature. The temperature and pressure are held for a set length of time, and then the gas is cooled and vented.
The HIP treatment process is used to produce near-net shaped components, reducing or eliminating the need for subsequent machining operations. Further, by precise control of the temperature, pressure and time of a HIP treatment schedule a particular microstructure for the treated part can be obtained.
All casting processes must deal with problems that the wrought processes do not encounter. Major among those are porosity and shrinkage that are minimized by elaborate gating techniques and other methods that increase cost and sometimes lower yield. However, the ability to produce a near-net or net shape is the motivating factor. In some cases, it is more cost effective to intentionally cast the part not using elaborate and costly gating techniques and HIP treat the part to eliminate the sub-surface porosity. The surface of the part is then machined until the dense substrate is reached.
U.S. Pat. No. 5,156,321, issued to Liburdi et al and U.S. Pat. No. 5,071,054, issued to Dzugan et al. are examples of methods that employ the HIP treatment process. Liburdi et al. discloses a technique to repair or join sections of a superalloy article. A powder matching the superalloy composition is sintered in its solid state to form a porous structure in an area to be repaired or joined. A layer of matching powder, modified to incorporate melting point depressants, is added to the surface of the sintered region. Liburdi discloses that the joint is raised to a temperature where the modified layer melts while the sintered layer and base metal remain solid. The modified material flows into the sintered layer by capillary action resulting in a dense joint with properties approaching those of the base metal. This reference discloses that HIPing can be used as part of the heat treatment to close any minor interior defects. Dzugan et al. discloses fabricating a superalloy article by casting, and then refurbishing primary defects in the surface of the cast piece. The defects are removed by grinding. The affected portions of the surface are first filled with a material that is the same composition as the cast article. Then, a cladding powder is applied to the surface through the use of a binder coat to obtain a smooth surface. The article is then heated to melt the cladding powder, and then cooled to solidify. Finally, the article is HIPed to achieve final closure of the surface defects.
Metal alloy components, such as gas turbine parts such as blades and vanes, are often damaged during use. During operation, gas turbine parts are subjected to considerable degradation from high pressure and centrifugal force in a hot corrosive atmosphere. The gas turbine parts also sustain considerable damage due to impacts from foreign particles. This degradation results in a limited service life for these parts. Since they are costly to produce, various repair methods are employed to refurbish damaged gas turbine blades and vanes.
Some examples of methods employed to repair gas turbine blades and vanes include U.S. Pat. No. 4,291,448, issued to Cretella et al.; U.S. Pat. No. 4,028,787, issued to Cretella et al.; U.S. Pat. No. 4,866,828, issued to Fraser; and U.S. Pat. No. 4,837,389, issued to Shankar et al.
Cretella '448 discloses a process to restore turbine blade shrouds that have lost their original dimensions due to wear while in service. This reference discloses using the known process of TIG welding worn portions of a part with a weld wire of similar chemistry as the part substrate, followed by finish grinding. The part is then plasma sprayed with a material of similar chemistry to a net shape requiring little or no finishing. The part is then sintered in an argon atmosphere. The plasma spray process used in accordance with Cretella '448 results in a coating porosity of about 5.0%. Even after sintering the coating remains attached to the substrate and weld material only be a mechanical bond at an interface bonding layer making the finished piece prone to chipping and flaking.
Cretella '787 discloses a process for restoring turbine vanes that have lost their original dimensions due to wear while in service. Again, a conventional plasma spray process is used to build up worn areas of the vane before performing a sintering operation in a vacuum or hydrogen furnace. The porosity of the coating, and the interface bonding layer, results in a structure that is prone to chipping and flaking.
Fraser discloses a process to repair steam turbine blades or vanes that utilize some method of connecting them together (i.e. lacing wire). In accordance with the method disclosed by Fraser, the area of a part that has been distressed is removed and a new piece of like metal is welded to the part. The lacing holes of the part are plug welded. The part is then subjected to hot striking to return it to its original contour, and the lacing holes are re-drilled.
Shankar et al. disclose a process for repairing gas turbine blades that are distressed due to engine operation. A low-pressure plasma spray coating is applied to the vanes and the part is re-contoured by grinding. A coating of aluminum is then applied using a diffusion coating process. Again, the conventional low-pressure plasma spray process forms a mechanical bond at an interface boundary between the coating and the substrate, resulting in a structure that is prone to failure due to chipping and flaking.
Other examples of methods for repairing or improving the characteristics of turbine engine airfoil parts include U.S. Pat. No. 5,451,142 issued to Cetel et al.; U.S. Pat. No. 4,921,405, issued to Wilson; U.S. Pat. No. 4,145,481 issued to Gupta et al.; and U.S. Pat. No. 5,732,467 issued to White et al.
Cetel discloses a turbine engine blade having a blade root with a surface having a thin zone of fine grains. A plasma spray technique is used to form a thin layer of material on the root or fir tree portion of the blade. The blade is then HIPed. After the HIP process, the blade is solution heat treated and then machined. This reference is directed to a process for modifying the root section of a turbine blade to improve the mechanical properties of this area of the part. The root section is serrated and is attached to the disc by inserting the root serrations into matching serrations of the disc. The blade is normally produced, as relating to chemistry and microstructure, to maximize the creep rupture and high cycle fatigue properties of the airfoil which is exposed to the hot gas path. The root section of the part thus has those same properties as the airfoil section. However, the root section of the blade is exposed to stress of a type different than the airfoil section, usually referred to as low cycle fatigue. The root section experiences colder operating temperatures than the airfoil section and is not directly in the path of the hot gasses that flow through the engine. Also, the root section is subjected to metal to metal stress during rotation resulting in low cycle fatigue cracking. Cetal is concerned with treating only the fir tree or root portion of the blade to improve its mechanical properties. The root portion or a new or refurbished blade is treated with a plasma spray process, HIPing, and a heat treatment and then machined. The blade is machined to remove material from a high stress portion of the blade root. The material removed by the machining operation is replaced by a zone of fine grains by a plasma spray technique. The part is processed through a HIP cycle to densify the deposit, and then a heat treatment cycle to enhance its properties. Finally, the root is machined back to the desired blueprint dimensions and the part returned to service.
Wilson discloses a turbine engine blade having a single crystal body having an airfoil section and an attachment or root section. A layer of polycrystalline superalloy is applied to the attachment section, preferably by plasma spraying. The coated blade is HIPed and then solution heat-treated to optimize the polycrystalline microstructure.
Grupta discloses a process for producing high temperature corrosion resistant metal articles. A ductile metallic overlay is formed on the surface of an article substrate, and an outer layer is applied over the overlay. The article is then subjected to a HIP treatment to eliminate porosity and create an inter-diffusion between the outer layer the overlay and the substrate.
None of these prior attempts provide for the effective and efficient restoration of the critical airfoil dimensions of a gas turbine engine airfoil part. Typically, an airfoil part will have to be discarded after it has gone through a certain number of repair cycles. The stripping of the protective coating on the part during the repair process is a major contributing factor resulting in the discarding of the part. After a number of repair cycles the part simply does not have the minimum dimensional characteristics necessary for it to perform it intended function. Therefore, there is a need for a method for repairing gas turbine engine airfoil parts that effectively and efficiently restores the critical airfoil dimensions of the part.
Turbine engine airfoil parts, such as vanes, are manufactured to precise tolerances that determine the airflow characteristics for the part. The class of a turbine vane is the angular relationship between the airfoil section and the inner and outer buttresses of the vane. This angular relationship has a direct bearing on the angle of attack of the airfoil section during the operation of the gas turbine engine. Over time, the angular relationship between the airfoil section and the inner and outer buttresses of the vane may become altered due to, for example, deformation of the airfoil section from engine operation and repair processes and the like. Or, the particular angular relationship of the airfoil section and the inner and outer buttresses as originally manufactured may need to be changed to improve engine performance. In any event, there is a need for a method of restoring or reclassifying a gas turbine engine airfoil part.